Launch vehicle and system and method for economically efficient launch thereof

ABSTRACT

The present disclosure relates to a launch system, a launch vehicle for use with the launch system, and methods of launching a payload utilizing the launch vehicle and/or the launch system. The disclosure can provide for delivery of the payload at a terrestrial location, an Earth orbital location, or an extraorbital location. The launch vehicle can comprise a payload, a propellant tank, an electrical heater wherein propellant, such as a light gas (e.g., hydrogen) is electrically heated to significantly high temperatures, an exhaust nozzle from which the heated propellant expands to provide an exhaust velocity of, for example, 7-16 km/sec, and sliding electrical contacts in electrical connection with the electrical heater. The launch vehicle can be utilized with the launch system, which can further comprise a launch tube formed of concentric electrically conductive tubes, as well as an electrical energy source, such as a battery bank and associated inductor.

CROSS-REFERENCE TO RELATED APPLICATIONS

The present application is a continuation of U.S. patent applicationSer. No. 14/211,698, filed Mar. 14, 2014, and claims priority to U.S.Provisional Patent Application No. 61/799,931, filed Mar. 15, 2013, thedisclosures of which are incorporated herein by reference in theirentireties.

FIELD OF THE DISCLOSURE

The present disclosure relates to systems, methods, and apparatuses forlaunch of a payload. More particularly, the payload may be intended forspace launch or delivery to a terrestrial location, and the presentdisclosure can provide for acceleration of the payload from a launchtube to its desired location.

BACKGROUND

Many means are known for accelerating an object. Relatively smallprojectiles are efficiently accelerated via controlled explosivecharges, such as with gunpowder. As the mass of the object beingaccelerated increases, however, the explosive force required greatlyincreases. For example, chemical combustion rockets are presently theonly means that have been shown to be effective for launching payloadsinto space. Although much work has been done in attempting to developalternative technologies for rapid acceleration of large payloads, suchas electromagnetic launchers and plasma arc acceleration, no alternativetechnology to date has been proven useful and reliable, particularly inthe launch of space vehicles.

In relation to space launch, while rocket propulsion is a long proventechnology, reliance solely on rocket launch is problematic in that itremains expensive, dangerous, and is dominated by government funding.Such problems are illustrated by the retirement and lack of suitablereplacement for the National Aeronautic and Space Administration's SpaceShuttle program. The lead time for a new space launch using rocketpropulsion is typically three to ten years. Space launches areinfrequent, typically occurring less than once a year to a few times peryear per customer. This has hampered advancements in certaintechnologies, such as communications. For example, satellitetechnologies have been slow and expensive to develop and are oftenoutdated quickly after launch and satellite placement. These factors andattendant continued government involvement have locked in high costs andlow profits. In particular, it is widely understood that present rocketlaunch technology can cost greater than $20,000 per kilogram of materialfor placement in an earth orbit.

Many types of gun launch systems have been proposed. A taxonomy of thevarious types is shown below.

All of the foregoing gun launch approaches share a common feature inthat they impose unacceptably great acceleration forces on the payload.Accelerations are typically tens of thousands of G's. These imposeenormous challenges in designing payloads that can survive the launchand still accomplish complex tasks after launch. Despite thesechallenges, perceived payoffs were so high that the US governmentinvested hundreds of millions in R&D on all the various types of gunlaunchers in the 1970's through the early 1990's.

A few commercial entities have attempted to enter the space launchmarket; however efficient and reliable launch means are yet to beproven. Moreover, the persistently high costs of space launch meangovernment spending will continue to be an important factor in spacelaunch technologies, and profitability will continue to remain low.Accordingly, there remains a need in the art for systems, methods, andapparatuses for reliable and efficient launch of projectiles, includingspace vehicles.

SUMMARY OF THE DISCLOSURE

The present disclosure relates to launch vehicles, launch systems, andmethods of launching a payload. The disclosure provides for delivery ofthe payload to any desired location on Earth, in earth orbit, or inspace generally. The payload can comprise a variety of objects,including satellites, raw materials or resources, ballistics, and thelike. The payload further can include human passengers.

In one aspect, the present disclosure provides a launch vehicle.Preferably, the launch vehicle is adapted for high velocity delivery ofa payload. In certain embodiments, the launch vehicle can comprising thefollowing: a payload container; a propellant tank containing apropellant; an electrical heater in fluid connection with the propellanttank and adapted for electrical heating of the propellant to form anexiting exhaust; and one or more electrical contacts adapted fordirecting flow of electrical current through the electrical heater. Infurther embodiments, the launch vehicle also can comprise an expansionnozzle in fluid communication with the exiting exhaust from theelectrical heater. In some embodiments, the electrical heater can be aresistive heater. In particular, the resistive heater can comprise anelectrically heated porous cylinder inside a containment vessel.Specifically, the electrically heated porous cylinder can comprisetungsten walls. In further embodiments, the electrical heater can be anarc heater. Preferably, the arc can be a swirl stabilized vortex arc.Further, the arc heater can comprise a swirl chamber inside acontainment vessel. In some embodiments, the arc heater can comprisecoaxial electrical terminals spaced apart by the swirl chamber. In otherembodiments, the electrical contacts can be sliding electrical contacts.

In a further aspect, the present disclosure provides a launch system. Invarious embodiments, the launch system can comprise the following: alaunch vehicle as described herein; a launch tube comprising two or moreconcentric, electrically conductive tubes separated by an insulator, thelaunch tube being adapted for propulsion of the launch vehicletherethrough; and an electrical energy source. In some embodiments, theelectrical energy source can comprise a battery bank. In furtherembodiments, the electrical energy source further can comprise aninductor.

In still another aspect, the present disclosure provides a method forlaunching a payload. In certain embodiments, the method can comprise thefollowing steps: providing the payload in a payload container of alaunch vehicle included in the launch system as described herein; andelectrically heating the propellant in the electrical heater to form theexiting exhaust at a velocity sufficient to propel the payload out ofthe launch tube.

In further embodiments, the present disclosure can be characterized by anumber of different embodiments. In particular, the presently disclosedlaunch vehicle, launch system, and method for launching a payload can bedefined by one or more of the following statements.

The disclosure encompasses a launcher in which a launch package isaccelerated from a launch tube using material exhausted from the launchpackage wherein the exhaust is formed by the heating of a low atomicweight element contained within (or in front of and in contact with) thelaunch package, wherein the energy used for such heating is providedelectrically to the launch package from the launch tube walls.

The launcher is located on the earth.

The launcher is located in free space.

The launcher is located on another celestial body.

The exit velocity is in the range of 2,000-50,000 m/sec.

The exit velocity is 4,000-30,000 m/sec.

The exit velocity is 6,000-15,000 msec.

The exit velocity is 8,000-12,000 m/sec.

The launch package is first accelerated to an initial velocity of100-5,000 msec using a single stage light gas gun.

The velocity is 500-3,000 msec.

The velocity is 1500-2500 msec.

The light gas is preheated.

The gas is electrically heated.

The electrical heating is derived from the same energy supply as thelauncher.

The tube is constructed with two concentric conductors one inside theother with minimum thickness of insulation between them so as tominimize the volume of the magnetic field “charged up” by the high drivecurrents going down and back up the tube to drive a launch packageheater. This magnetic field energy has several deleterious effects. Anundesired magnetic field requires energy which does not help propel thelaunch package. The undesired magnetic energy may be dischargedimmediately upon launch and can cause catastrophic damage if the energylevel is not minimized (as is achieved by the present disclosure). Thishas caused prior art launchers, such as railguns, to fail to achievehigh velocities. The magnetic field produces high mechanical forces andstresses which cause breakdown or higher cost or wear. This also hascaused prior art launchers, such as railguns, to fail to achieve highvelocities. The magnetic field produces high induced voltages which cancause arcing in places where it is not wanted and can cause wear orcatastrophic damage or failure. The magnetic field produces a force onany arc formed between the launch tube conductors and the slidingcontacts that transmit electrical current and energy to the launchpackage. Usually these forces cause the arc to blow forward at up to thespeed of light. The electrical current is then diverted away from whereit is wanted and needed to propel the launch package to a different areawhere it is not wanted, causing loss of energy, retarding of the launchpackage, wear, and/or catastrophic damage and/or failure. This furtherhas caused prior art launchers, such as railguns, to fail to achievehigh velocities.

The present disclosure provides an “electroantimagnetic” launcher. Priorart launchers, such as railguns and coil guns, actively inducedformation of magnetic forces to propel the launch package. The requiredgeneration of very high magnetic fields led to the aforementioneddeleterious effects.

The present disclosure minimizes magnetic fields.

The present disclosure uses the force of hot expanding gas formed byelectrical heating facilitated by tube-conducted electrical energy.

The present disclosure utilizes expansion of hot, light gas.

Hydrogen can be heated above 5,000 K and result in an expelled gasconsisting of individual atoms of hydrogen.

Heating can be up to 100,000 K, which can result in the exhaust velocitybeing 77,000 m/sec.

Launch velocity can be as great as 150,000 m/sec or about two times theexhaust velocity.

Maximum velocities can be limited to use in space since the practicallimit for launch from the surface of the Earth is about 100,000 msec dueto aerodynamic forces at that speed, which can reach about 1,000,000PSI.

Velocities with Earth launch can practically reach about 50,000 msecwhich produces 250,000 PSI, which can be mitigated using transpirationcooled metal nosetips.

Velocities with Earth launch can practically reach about 18,000 msec,which produces about 30,000 PSI, which can be mitigated by ablatingcarbon nosetips. Another limit is imposed by launcher length when peopleare launched.

Velocities for launchers ferrying human passengers can be limited to anacceleration of no more than about 20 G's. Using a launch tube with alength up to about 1,000 km can safely lead to velocities of up to about20,000 msec.

The light gas is heated in an electrical heater.

The heater is a resistive heater.

The heater is an arc heater.

The heating element is a transpiration tube element.

The heater wall is cooled by transpiration.

The light gas is seeded with an ionizable element such as cesium orrubidium or potassium or sodium or lithium to promote arc stability andconductivity and ionization.

The conductive sliding contacts make mechanical sliding contact with thetube walls with a low voltage drop.

The conductive sliding contacts make an arcing sliding contact with thetube walls with a minimized voltage drop.

The arc is contained via mechanical containment using a slidinginsulating perimeter.

The arc is contained via magnetic forces.

The magnetic forces are generated by the current transferring throughthe contact via a specially shaped current loop.

The magnetic forces are generated by self contained power source on thelaunch package.

The magnetic forces are generated by a magnet.

The magnet is a superconducting magnet.

The sliding contact is cooled by transpiring fluid.

The sliding contact is cooled by material in the sliding contact meltingor vaporizing.

The transpiring fluid is conductive.

The fluid is a low melting metal having a low ionization potential.

The metal is cesium, aluminum, lithium or analogous low melting softmetals with low ionization potentials.

The insulating perimeter is transpiration cooled.

The transpiration fluid is an insulating material such as hydrogen,sulfur hexafluoride, or other liquid or gas.

At least a portion of the sliding contact can be adapted to exhibit oneor more state transitions.

The sliding contact can define a sliding solid-solid interaction with asolid tube wall that transitions to a liquid-solid interaction when atleast a portion of the sliding contact that interacts with the tube walltransitions to a liquid metal melt.

The state transition occurs at a launcher velocity of about 1000 toabout 2000 msec.

The state transitions to an arcing contact at a velocity of about 1500to about 3000 msec.

Arc voltage can be about 100 to abut 300 V.

The arc is stably positioned at the contact and does not substantiallymove outside of the desired contact region.

The conductive tube walls have slotted tracks of varying numbers andgeometries for the sliding contacts to make contact with conductivestrips and which prevent arcing between the tube conductors and whichalso serve to align the launch package and keep it from rotating in thetube.

The conductive strips define longitudinal tracks extending at least aportion of the length of the conductive tube.

The conductive strips are coaxial with the conductive tube.

The conductive tube walls have layers of different materials.

The predominant material is steel or aluminum.

The innermost layer is a high temperature wear resistant conductivematerial such as tungsten or rhenium or hardened copper.

An interlayer of material is between the predominant outer layer and theinnermost layer.

The interlayer is copper or molybdenum.

A majority of the inner surface of the conductive tube walls is coatedwith an insulator.

Substantially all of the inner surface of the conductive tube wallsapart from the sections defined by the conductive strips is coated withan insulator.

The insulator on the inner wall of the conductive is a ceramic or acomposite.

There is one current outbound path in series with one return currentpath.

There are multiple current outbound paths in parallel, and all in serieswith multiple return current paths.

The launcher inductance is lowered proportionately to the number ofparallel current paths.

The lower inductance lowers the magnetic field energy and thus thedetrimental effects of the magnetic field.

The slotted track insulators are transpiration cooled.

The conductive strips are transpiration cooled.

The transpiration fluid is a conductive material.

The conductive strips are conductively and/or convectively cooled.

The fluid is a low melting metal having a low ionization potential.

The metal is cesium, aluminum, lithium or analogous low melting softmetals with low ionization potentials.

The propellant tank has an outer diameter that is substantiallyidentical to the inner diameter of the launch tube.

The propellant tank includes sliding contact strips on at least aportion of its outer surface.

The sliding contact strips vaporize as the velocity of the launchvehicle increases.

The vaporized strips provide a low drag gas bearing to minimizefrictional drag.

The strips produce a vapor that is insulating so that it inhibits ratherthan promotes any arcing.

The sliding contact strips comprise pores filled with liquid sulfurhexafluoride.

A device can be inserted into the conductive launch tube for inspection,alignment, and repair.

The launch tube is aligned by active alignment devices.

The launch tube is substantially horizontal except near the exit endwhere it curves upward.

The launch tube follows the curvature of the Earth.

The launch tube is at a constant slope angle.

The tube bed is graded to the tube constant slope angle.

The launch tube is evacuated and backfilled with a low pressure of lightgas to minimize aerodynamic drag during acceleration while providingincreased resistance to arc breakdown ahead of the launch package.

The launch tube is evacuated and a high speed pulse of gas is introducedtime sequentially along the tube via transpiration to coat the wallswith a layer of gas which insulates the walls but does not have time toexpand into the majority of the tube diameter and thus increaseaerodynamic drag.

An initial section of the launch tube is not electrical conductive.

The launch vehicle is accelerated through the initial section of thelaunch tube utilizing an expanding gas that is not electrically heated.

The launch tube exit is sealed with an exit device to prevent airingress until the launch package arrives.

The exit device is a high speed mechanical shutter.

The exit device is one or a series of aerodynamic curtains.

The exit device is a thin membrane or membranes which the launch packageflies through.

The exit device is a thin membrane or membranes, with one or severalsmall explosive charges which destroy the launch package if notdetonated prior to launch package arrival and which allow the launchpackage to pass if detonated prior to launch package arrival.

The launch tube is moveable.

The launch tube is moveable in one dimension to change launch elevationor launch azimuth.

The launch tube is moveable in two dimensions to enable change in bothelevation and azimuth.

The launch tube is mounted on a moveable vehicle such as a ship or asubmarine.

The launch tube is installed in a slanted tunnel underground.

The launch tube is installed on naturally sloping ground.

The launch package has inertial sensors and actuators which activelymaintain its alignment and orientation while being accelerated in thelaunch tube.

The launch package is monitored during the launch acceleration intervalfor integrity and nominal performance.

Emergency procedures can be implemented based on monitoring results tooptimize the launch and to protect the launch tube.

The launch can be aborted by destroying the launch package immediatelyor shortly after its exit from the launch tube.

The launch package is separated during or immediately after launch intoa discarded portion and a flyout payload portion.

The flyout payload has a heat shield with a transpiration cooled nosetipto maintain the nosetip integrity, shape, sharpness, low drag, and lowpressure moment during exit from the atmosphere.

The flyout payload has a small positive stability, neutral stability, ora negative aerodynamic stability based on its center of pressurelocation relative to its center of mass location.

The flyout payload can maneuver at high lateral acceleration levels tooptimize flight path through the atmosphere and change launch azimuth.

The flyout payload has a high lift to drag ratio.

The payload has a lifting body design.

The flyout payload has aerodynamic control surfaces with very high speedresponse and low drag.

The control surfaces are base split flaps.

The control surfaces are actuated with piezoelectric actuators.

The flyout payload is an orbital satellite.

The flyout payload is a suborbital payload.

The satellite is a communications satellite or a sensor satellite orresupply vehicle or a weapon.

The payload is a commercial package to be delivered rapidly to longdistances.

The payload is a sensor payload or a UAV or other unmanned vehicle.

The payload is a weapon.

The payload contains subparts that are dispersed before impact.

The payload remains intact until impact.

Multiple payloads impact at or near the same location for deeppenetration.

The satellite contains an inflatable solar array for power.

The inflatable structure hardens to rigidity after deployment.

The satellite contains an inflatable magnet array to effect attitudecontrol in orbit.

The inflatable structure hardens to rigidity after deployment.

The satellite contains an inflatable antenna array to effectcommunications in orbit.

The inflatable structure hardens to rigidity after deployment.

The satellite contains inflatable structures to effect missions inorbit.

The inflatable structures harden to rigidity after deployment.

The design lifetime of the satellite is less than 10 years, or less than5 years, or less than 2 years, or less than 1 year.

The satellite orbital altitude is such that the orbital lifetime due toaerodynamic drag is less than 5 years, or less than 2 years, or lessthan 1 year, or less than 6 months, or less than 3 months, or less than1 month.

The satellite achieves longer orbital lifetime through magnetic thrustagainst the Earth's magnetic field using an inflatable magnetic array.

The satellite achieves longer orbital lifetime through pressure inducedby sunlight and solar wind on an inflatable solar sail.

The satellite achieves longer orbital lifetime throughmagnetohydrodynamic (MHD) propulsion against ionized upper atmospheremolecules.

The payload cost is reduced through using commercial grade parts withhigh initial failure rates and then iterating quickly through launch,fail, and redesign cycles to achieve higher and higher reliabilityquickly over time.

The launcher and thousands of payloads are designed simultaneously for asingle purpose.

The payloads are all communications satellites.

The satellites are radiofrequency communication satellites.

The satellites are optical communications satellites.

The payloads are reflective relays for millimeter waves or opticalbeams.

The payloads are nuclear waste containers.

The light gas propellant for the launcher and/or the single stage lightgas gun is hydrogen.

The light gas is heated to 1000 to 100,000 K, to 2000-50,000 K, to2500-20000 K, to 3000 to 15000 K, to 3500 to 10000 K, to 3500 to 5000 K.

The exhausted gas is molecular hydrogen (0.002 kg/mole).

The exhausted gas is atomic hydrogen (0.001 kg/mole).

The exhausted gas is hydrogen plasma (0.0005 kg/mole).

The exhaust device contains a nozzle throat transpiration cooled with alight gas (e.g., hydrogen).

The exhaust device contains a nozzle transpiration cooled with a lightgas (e.g., hydrogen).

The exhaust device contains a porous nozzle throat in which the poresare filled with a material that absorbs heat by melting and/orvaporization and/or disassociation (e.g., solid hydrogen or lithium orice).

The exhaust device contains a porous nozzle in which the pores arefilled with a material that absorbs heat by melting and/or vaporizationand/or disassociation (e.g., solid hydrogen or lithium or ice).

The flyout payload has a heat shield with a porous nosetip filled with amaterial (e.g., solid hydrogen or lithium or ice) that absorbs heat bymelting and/or vaporization and/or disassociation to maintain thenosetip integrity, shape, sharpness, low drag, and low pressure momentduring exit from the atmosphere.

The electrical energy for the launch system is supplied by a batterybank.

The batteries are lead acid batteries.

The batteries are automotive batteries.

An inductor is interposed between the battery bank and the launcher suchthat the battery bank charges the inductor and then the inductor isswitched over to and discharges into the launcher tube.

The discharge into the launcher is initiated by explosively actuatedswitches.

The discharge switching is accomplished with conventional switches withcapacitor mediated arcing control.

The inductor has a core composed of a high permeability material.

The core is designed for high discharge rates and low eddy currentlosses.

The inductor is actively cooled.

The core is actively cooled.

The conductors are actively cooled.

BRIEF DESCRIPTION OF THE FIGURES

Having thus described the disclosure in the foregoing general terms,reference will now be made to the accompanying drawings, which are notnecessarily drawn to scale, and wherein:

FIG. 1 is a side view of a launch system according to an exemplaryembodiment of the present disclosure wherein a launch vehicle ispositioned within an electrically conductive launch tube;

FIG. 2 is a schematic of a launch tube according to an exemplaryembodiment of the present disclosure in comparison with a launch tubefrom a typical prior art railgun;

FIG. 3 is a graph showing electrical to kinetic energy conversionefficiency as a function of launcher magnetic field intensity for alauncher according to an exemplary embodiment of the present disclosureand a known art electromagnetic launcher;

FIG. 4 is a graph showing electrical to kinetic energy conversionefficiency as a function of launcher magnetic field intensity for alauncher according to an exemplary embodiment of the present disclosureacross a preferred range of minimized magnetic field strength;

FIG. 5 is a graph showing efficiency as a function of inductance for alauncher according to an exemplary embodiment of the present disclosureand a known art electromagnetic launcher;

FIG. 6 is a graph showing inductance per unit length versus launchergeometry for a launcher according to an exemplary embodiment of thepresent disclosure;

FIG. 7 is a graph showing efficiency versus launcher geometry for alauncher according to an exemplary embodiment of the present disclosure;

FIG. 8 illustrates sliding contact according to an exemplary embodimentof the present disclosure;

FIG. 9 is a rear view of a launch system according to an exemplaryembodiment of the present disclosure wherein a launch vehicle ispositioned within an electrically conductive launch tube;

FIG. 10 is schematic of a launch system according to exemplaryembodiment of the present disclosure showing a launch tube in connectionwith an electrical energy source;

FIG. 11 is a cross-section of an electrical heater according to anexemplary embodiment of the present disclosure comprising elementsuseful for resistive heating;

FIG. 12 is a cross-section of an electrical heater according to anexemplary embodiment of the present disclosure comprising elementsuseful for swirl stabilized vortex arc heating;

FIG. 13 is a side view of a launch system according to an exemplaryembodiment of the present disclosure wherein a launch vehicle ispositioned within an electrically conductive launch tube that includesdifferential pressurization;

FIG. 14 is an illustration of a payload component of a launch vehicleaccording to an exemplary embodiment of the present disclosure showingan external view of the payload component and an internal view of thepayload component revealing various elements of the exemplifiedembodiment, the payload component being in an atmospheric transitconfiguration; and

FIG. 15 is the payload component of a launch vehicle according to anexemplary embodiment of the present disclosure shown in FIG. 6, whereinthe payload component is in an on-orbit deployed configuration.

DETAILED DESCRIPTION OF THE DISCLOSED EMBODIMENTS

The present disclosure will now be described more fully hereinafter withreference to exemplary embodiments thereof. These exemplary embodimentsare described so that this disclosure will be thorough and complete, andwill fully convey the scope of the disclosure to those skilled in theart. Indeed, the present disclosure may be embodied in many differentforms and should not be construed as limited to the embodiments setforth herein; rather, these embodiments are provided so that thisdisclosure will satisfy applicable legal requirements. As used in thespecification, and in the appended claims, the singular forms “a”, “an”,“the”, include plural referents unless the context clearly dictatesotherwise.

The present disclosure provides means for rapid acceleration of aprojectile. In a particular embodiment, the projectile can comprise allor part of a space launcher. As such, the disclosure may focus on thisembodiment for simplicity in describing the several features of thepresent subject matter. Nevertheless, the disclosed subject matter isnot intended to be limited to space launch or to further specificembodiments discussed herein. Rather, any disclosure in relation to aspecific embodiment is intended to be exemplary of the subject matter soas to provide a description sufficient to extend the exemplarydiscussion to further embodiments.

The present disclosure provides a launch system, one or more apparatusesthat can be utilized in the launch system, and one or more methods oflaunching an object, in particular to a high velocity at rapidacceleration. In certain embodiments, these and further aspects of thedisclosure can be achieved though use of an electroantimagnetic (“EAM”)launcher that utilizes low atomic weight elements as a propulsionpropellant.

Electromagnetic (“EM”) launchers have been under active development bythe U.S. government and other entities for approximately 30 years. EMlaunchers (e.g., railguns) rely upon induced magnetic fields created byelectrical current flowing down two parallel rails and through a launchapparatus that is accelerated by the electromagnetic effect. Similarly,augmented EM launchers can utilize a driving current that is channeledthrough additional pairs of parallel conductors that are arranged so asto increase or augment the magnetic field experienced by the launchapparatus. Research to date, however, indicates that EM launchers sufferfrom severe problems that prevent them from achieving the conditionsnecessary for space launch. At high velocities, magnetic effects absorbenormous power and energy. This can effectively destroy the ability toachieve stable electrical conduction between a moving launch package andstationary feed conductors. Rather than using electrical energy tocreate magnetic forces to propel the launch package, the EAM launcheraccording to the present disclosure can substantially minimize inducedmagnetic forces and thus likewise minimize the energy drain anddistorting effects thereof.

A launch system according to the present disclosure can comprise anumber of components that may independently provide useful improvementsover known technologies as well as in a number of combinations of thedisclosed components. For example, in some embodiments, a launch systemaccording to the present disclosure can comprise a launch tube inelectrical connection with an electrical energy source so as to provideelectrical energy to a launcher that may comprise one or more of apropellant source, an electrical heater for heating the propellant fromthe propellant source, sliding contacts in electrical contact with theelectrical heater and in electrical contact with the launch tube, anexpansion nozzle in fluid communication with the electrical heater andadapted for discharge of one or more heated propellant or a componentthereof, and a payload in mechanical connection with one or more of thefurther components of the launcher.

A side view of components of a launch system 20 according to oneembodiment of the present disclosure is shown in FIG. 1. As illustratedtherein, a launcher 200 is positioned within a launch tube 100. Thelaunch tube 100 can comprise a plurality of concentric, conductive tubesthat can be separated by one or more insulating layers. Alternatively,the launch tube can comprise a single, multi-layer tube comprising aplurality of conductive layers separated by one or more insulatinglayers. As shown in the embodiment of FIG. 1, the launch tube 100 cancomprise an outer conductive tube 110 spaced apart from an innerconductive tube 130 and separated by an insulator 120. The outerconductive tube 110 and the inner conductive tuber 130 can be formed ofany suitable, conductive material, such as a metal or metal alloy. Insome embodiments, the conductive tube walls can comprise layers of twoor more different materials. As exemplary embodiments, one or both ofthe inner conductive tube and the outer conductive tube can comprisesteel, aluminum, or an aluminum alloy. In preferred embodiments, theinnermost layer of one or both of the inner conductive tube and theouter conductive tube can comprise a high temperature wear resistantconductive material such as tungsten, rhenium, or hardened copper. Assuch, one or both of the inner and outer conductive tubes can comprisean outer layer formed of steel or aluminum and an inner layer formed ofthe high temperature wear resistant conductive material (wherein innerreferences proximity to the interior of the tube and outer referencesproximity to the exterior of the tube). In some embodiments, aninterlayer material can be formed between the predominant outer layerand the inner layer. For example, an interlayer can comprise copper ormolybdenum.

The insulator 120 can comprise any material effective to substantiallyprevent flow of electrical current between the two conductive tubes. Inpreferred embodiments, the thickness of the insulator 120 and the totalspace between the two conductive tubes can be minimized. Suchminimization can be useful to minimize the volume of a magnetic fieldformed by the electric current flow through the conductive tubes. Forexample, in various embodiments, electric current passes through one ofthe inner and outer conductive tubes, through an electrical heater, asfurther discussed below, and back along the other of the inner and outerconductive tubes. The thickness of the insulator and/or the total spacebetween the outer conductive tube and the inner conductive tube can beabout 0.5 cm to about 30 cm, about 1 cm to about 20 cm, about 1.5 cm toabout 15 cm, or about 2 cm to about 10 cm.

The advantages of the present EAM launcher are evident in comparison torailgun technology. The geometry of the launch tube in particular forthe present EAM launcher can significantly reduce or eliminate theadverse effects of large magnetic fields. The key concept driving theknown art of EM launchers, such as railguns, has been to maximize themagnetic forces pushing the projectile. In attempts to achieve thegreatest launch force, very large magnetic fields are used with EMlaunchers. This involves a significantly large electrical currentrequirement and results excessive mechanical pressures from the intensemagnetic fields. It also results in very large resistive losses and arclosses. As an example, a typical railgun designed to launch a one tonpayload uses 10 million to about 50 million amps of electrical currentand magnetic fields of about 10 to about 25 Tesla, with resultingpressures of about 15,000 to about 100,000 PSI. Losses in the systemlead to a railgun having roughly 10% efficiency when evaluated atorbital velocities, and this only if one can effectively maintain thestructural integrity of the plasma armature. That feat has heretoforenever been accomplished, despite government research and developmentinvestments on the order of about $1B. The highest recorded, repeatablevelocity ever achieved with railgun technology is about 6,000 meters persecond, and the highest efficiencies of such high velocity launches wasonly a few percent.

The EAM launcher according to various embodiments of the presentdisclosure can require application of an electrical current of about 0.2to about 2.0 million amps—i.e., a 5 to 200 fold reduction in comparisonto typical EM launcher technology. This is particularly relevant sincelosses and pressures scale as the square of applied electrical currentand thus are reduced 25 to 40,000 fold. With that basis, an EAM launcheraccording to embodiments of the present disclosure can provide dramaticincreases in efficiency as compared to EM launcher technology—as much aseven a 10,000 fold improvement.

The advantages of the presently disclosed EAM launcher are furtherevident in the various appended figures. For example, FIG. 2 shows aside-by-side comparison of an exemplary embodiment of a launch tube foran EAM launcher and a launch tube for a typical, known EM railgun.Distinct differences are readily apparent. For example, in the EAMlauncher, the conductors can completely encircle the launch tubewhereas, in the EM launcher, the conductors form significantly less than50% of the surface area of the launch tube internal wall. Rather,insulators form the significant majority of the inner launch tube wallsurface. The effect of the very large space separating the electricalconductors in the EM launch tube is a large volume, highly intensemagnetic field, and this is desired in typical railguns to drive theprojectile. On the contrary, in the present EAM launcher, the concentricconductors are separated by only a small distance. As such, the totalvolume of space between the conductors is minimized and effectivelylimits that intensity of the induced magnetic field. The significantdifference in magnetic field strength is illustrated in FIG. 3.

The electrical to kinetic energy conversion efficiency of a highvelocity launcher can be a function of, at least in part, the strengthof the magnetic field within the launch tube. Typical, known EMlaunchers require magnetic fields in the range of 10 Tesla to greaterthan 20 Tesla. As illustrated in FIG. 3, the electrical to kineticenergy conversion efficiency of the EM launcher increases as themagnetic field strength increases; however, the highest efficiency (at afield strength of about 25 Tesla is only in the range of about 12%(which is an increase from the lowest efficiency of about 1%). Thecompletely opposite effect is seen with the EAM launcher of the presentdisclosure. Specifically, the efficiency of the EAM launcher dropsdramatically as the magnetic force increases. The present EAM launcher,however, surprisingly can achieve an electrical to kinetic energyconversion efficiency of greater than 51% in various embodiments whenmagnetic field strength is substantially 0 Tesla. Accordingly, incertain embodiments of the present disclosure, the EAM launcher, andparticularly the launch tube, can be adapted to transmit necessaryranges of electrical current while generating or inducing a magneticfield at a strength of less than about 1.25 Tesla, less than about 1Tesla, less than about 0.5 Tesla, less than about 0.25 Tesla, or lessthan about 0.2 Tesla. In some embodiment, magnetic field can be limitedto a strength of about 0.2 Tesla to about 1.2 Tesla. This is graphicallyillustrated in FIG. 4. In some embodiments, the magnetic field strengthcan be sufficiently minimized through appropriate minimization of theinsulator volume between the concentric conductive tubes. In otherembodiments, however, further means for limiting, reducing, oreliminating any magnetic field can be utilized.

Efficiency for an EAM launcher according to the present disclosure canbe significantly improved over known art EM launcher technology inrelation to further properties. For example, FIG. 5 illustrates theadvantages of the presently disclosed EAM launcher over typical, priorart EM launchers in relation to inductance. As seen in FIG. 5, a typicalEM launcher seeks to increase inductance per unit length to increaseefficiency. On the other hand, efficiency for an EAM launcher accordingto the present disclosure can be maximized when inductance is minimized.

In order to minimize inductance according to certain embodiments of thepresent disclosure, it can be desirable to provide the launch tube forthe EAM launcher with a desired geometry. As shown in FIG. 6, inductanceper unit length can vary based upon the ratio of the outer radius to theinner radius for the launch tube (i.e., the ratio of the radius of theouter tube to the radius of the inner tube). Inductance can decrease asthe ratio decreases. Thus, it can be desirable to provide the launchtube with a low ratio—i.e., a minimized total thickness for any spaceexisting between the walls of the conductive launch tube. In certainembodiments, the ratio of outer to inner radius for a launch tubeaccording to the present disclosure can be less than 2, less than about1.5, less than about 1.25, less than about 1.15 or less than about 1.1.In further embodiments, the ratio of outer to inner radius can be about1.4 to about 10, about 1.5 to about 7.5, about 1.6 to about 5, about1.65 to about 4, or about 1.7 to about 2.5.

Launcher geometry also can affect efficiency of an EAM launcheraccording to the present disclosure. As seen in FIG. 7, the efficiencyof an EAM launcher can increase as the ratio of the outer to innerradius for the launch tube decreases. Thus, it further can be desirableto achieve the ratios already noted above. Specifically, it can bedesirable to approach a ratio that is as close to 1 as possible (limitedonly by the necessary thickness of the conductors for carrying currentand the thickness of insulators present to prevent high voltagebreakdown between the conductive tubes).

The walls of the outer and inner conductive tubes can have one or moreslotted tracks of varying geometries that are adapted for receiving oneor more sliding contacts. The cross-section of FIG. 1 passes through aslotted track 112 in the inner conductive tube 130 and the insulator 120(with a portion of the inner tube and insulator cut away to reveal thesliding contact in the slotted track). The slotted track 112 providesfor an electrical connection of the outer conductive tube 110 and anouter sliding contact 115. An inner sliding contact 135 is alsoillustrated in electrical connection with the inner conductive tube 130.The slotted tracks can be effective to facilitate proper contact betweenthe tubes and the sliding contacts, to prevent or substantially reducearcing between the tube conductors, and also serve to align the launcher200 and substantially prevent rotating thereof within the launch tube100. Aligning arms 113 and 133 can be in physical contact with thesliding contacts 115 and 135, respectively, and also be in physicalcontact with the electrical heater. The aligning arms preferablycomprise a high strength, rigid material, such as steel or a furthermetal alloy that can include an insulative layer, if desired.

The sliding contacts can be formed of a suitable conductive material andcan take on particular structures, as further discussed below.Preferably, the conductive sliding contacts are positioned so as todefine a mechanical sliding contact with the tube walls while exhibitingonly a low voltage drop. In some embodiments, the conductive slidingcontacts can define an arcing sliding contact with the conductive tubewalls with only a minimal voltage drop. The arc may be contained viamechanical containment, such as using a sliding insulating perimeter. Inother embodiments, the arc may be contained via magnetic forces, whichcan be generated by the current transferring from the contact. Inparticular, the contacts may define a current loop adapted to generatethe magnetic forces. In some embodiments, the magnetic forces can begenerated by a self contained power source that may be present on thelauncher. As an example, the magnetic forces can be generated by amagnet, which may be a superconducting magnet.

In some embodiments, the sliding contacts and/or the slotted trackinsulators can be cooled. Such cooling can be via conductive and/orconvective means. For example, the contacts may be cooled by a materialcomponent of the sliding contact that melts and/or vaporizes during use.In alternate embodiments, the sliding contact can be cooled by atranspiring fluid. The transpiring fluid can be conductive. Further, thefluid can be a low melting metal having a low ionization potential. Asnon-limiting examples, the metal can be cesium, aluminum, lithium, or ananalogous low melting soft metal with low ionization potential.Similarly, an insulating perimeter of the contacts and/or the alignmentarms can be transpiration cooled. For example, the transpiration fluidcan be an insulating material such as hydrogen, sulfur hexafluoride, ora like liquid or gas.

In certain embodiments, sliding contacts according to the presentdisclosure can be adapted to exhibit one or more state transitions. Forexample, the sliding contacts initially can define a sliding solid-solidinteraction with a solid tube wall. This solid-solid interaction cantransition to a liquid-solid interaction when at least a portion of thesliding contact that interacts with the tube wall is adapted totransition to a liquid metal melt. This can occur, for example, when thelaunch vehicle reaches a velocity of about 1000 to about 2000 m/sec.This can transition further, such as to an arcing contact. Suchtransition can occur, in certain embodiments, at a velocity of about1500 to about 3000 msec. A majority of electrical current transfer tothe sliding contact may occur during the arcing phase. The slidingcontacts can include mechanical, fluid dynamic, arc seeding, andelectromagnetic features to minimize the arc voltage and thus the energyloss at the contact. In some embodiments, arc voltage can be about 50 toabout 500 V. The arc preferably is stably positioned at the contact anddoes not substantially move outside of the desired contact region.

An exemplary sliding contact pad is shown in FIG. 8. The sliding contact500 includes an outer cooled rim 503 (e.g., transpiration cooled, suchas with hydrogen or with sulfur hexafluoride frozen into the pores). Inone embodiment, porous material with liquid SF₆ can be cooled withliquid hydrogen to freeze the SF6 in the pores. Alternatively, theliquid SF₆ can be sealed under pressure and released as the surfacemelts. Moving inward in FIG. 8, the sliding contact 500 further includesa magnetic rim 505, an ablative shield 507, an inner cooled rim 509, anablative rim 511, and the conductor 513.

The outer conductive tube and the inner conductive tube can define onecurrent outbound path in series with one return current path. In someembodiments, there can be multiple current outbound paths in parallel.If desired, all outbound paths can be in series with multiple returncurrent paths. The launcher inductance can be lowered proportionately tothe number of parallel current paths. Beneficially, the lower inductancecan lower the magnetic field energy and thus any undesired effects ofthe magnetic field.

The arrangement of the launch vehicle 200 and the launch tube 100 isfurther illustrated in FIG. 9, which shows an end view thereof. Again,the launch tube 100 includes an outer conductive tube 110 and in innerconductive tube 130 separated by an insulator 120. Sliding contacts(115, 135) are in electrical connection with the walls of the respectiveconductive tubes (110, 130). More particularly, outer sliding contact115 is in electrical connection with outer conductive strip 117, andinner sliding contact 135 is in electrical connection with innerconductive strip 137. The sliding contacts (115, 135) interconnect withthe electrical heater 220 via aligning arms (113, 133). The exhaustnozzle 210 is shown partially transparent to reveal some of the forwardcomponents of the launch vehicle 200.

In some embodiments, the launch tube can be aligned by active alignmentdevices. Further, the tube can be defined as being substantiallyhorizontal with the exception of a section defining and tube exit, wherethe tube may curve upward. The tube also can be characterized assubstantially following the curvature of the Earth. The tube can be at aconstant slope angle, and tube bed can be graded to the tube constantslope angle. Further, the launch tube can be installed on naturallysloping ground. Alternatively, the launch tube can be installed in aslanted tunnel underground. In certain embodiments, the launch tube canbe moveable. For example, the launch tube can be moveable in onedimension to change launch elevation or launch azimuth. Preferably, thelaunch tube can be moveable in two dimensions, as this can be beneficialto enable change in both elevation and azimuth. If desired, the launchtube can be mounted on a moveable vehicle such as a ship or a submarine.In certain embodiments, the launch tube can be defined by an initiallaunch section and a main launch section. The initial launch section canbe, for example, up to approximately 200 meters in length, and the mainlaunch section can be, for example, approximately 1,000 meters orgreater in length.

The launch tube specifically may be evacuated. Further, the launch tubemay be backfilled with a light gas, preferably at low pressure. This canbe beneficial to minimize aerodynamic drag during acceleration whileproviding increased resistance to arc breakdown ahead of the launchpackage. In particular embodiments, the launch tube can be evacuated,and a high speed pulse of gas can be introduced time sequentially alongthe tube so as to coat the tube walls with a layer of gas. This canfunction to insulate the tube walls but is present for a time sufficientto expand into the majority of the tube diameter and thus increaseaerodynamic drag. Such introduction of gas can be via transpirationthrough the tube walls.

The launch tube exit can be sealed with a device to substantially orcompletely prevent air ingress until the launch package arrives. Invarious embodiments, the exit seal can be, for example, a high speedmechanical shutter, one or a series of aerodynamic curtains, orrelatively a thin membrane or combination of multiple membranes throughwhich the launch package can safely fly. When the exit device is a thinmembrane or membranes, one or several small explosive charges may beprovided to destroy the membrane prior to arrival of the launch vehicleat the exit. Such charges can particularly function as a fail safemechanism. For example, the explosive charges may be used tointentionally damage a projectile prior to letting it leave the launchtube such that the projectile disintegrates almost immediately uponexiting the launch tube so as to abort a launch which does not meetspecified requirements.

Referring to FIG. 10, an electrical energy source 300 can be provided tosupply electrical energy to the launch tube 100, which comprises aninitial launch tube section 103, a main launch tube section 105, and alaunch tube exit 107. In certain embodiments, the electrical energysource 300 can comprise a battery bank. For example, a series of leadacid batteries (e.g., automotive batteries) may be used. Any furtherbattery or suite of batteries suitable for providing electrical energyon demand likewise may be used. In particular embodiments, an inductor350 can be interposed between the battery bank and the launch tube suchthat the battery bank charges the inductor while the inductor is in acharge state. Thereafter, the inductor 350 can be switched to adischarge state wherein the inductor discharges into the launch tube.The discharge into the launch tube may be initiated by explosivelyactuated switches. Alternatively, the discharge switching may comprisethe use of conventional switches with capacitor mediated arcing control.Preferably, the inductor can have a core comprising a high permeabilitymaterial. In particular, the core can be adapted for high dischargerates and low eddy current losses. Moreover, the inductor can beactively cooled, the core can be actively cooled, and/or the conductorscan be actively cooled.

Any power source adapted to provide about 200 to about 2000 kiloamps atabout 500 to about 5000 volts can be used as the electrical energysource according to the present disclosure. Non-limiting examples ofelectrical energy sources that may be used include capacitors, standardpower plant generators, rocket turbine driven turbogenerators, and thelike. In relation to cost and reliability, batteries (e.g., lead acidbatteries) driving an inductor as described above can be preferred.

The launch vehicle 200 is initially positioned inside the launch tube100 near the staging station 109 in the initial launch tube section 103.The launch system 20 can comprise additional elements as illustrated inFIG. 10, such as the payload preparation and launch operations building400. Briefly, in use, electrical energy is transferred from theelectrical energy source 300 via conduit 301 to the inductor 350 andthen through conduit 351 to the launch tube 100. The electrical energypasses through the conductive launch tube to the electrical heater 220via the sliding contacts (115, 135). The electrical energy specificallybetween the electrical heater 220 and the sliding contacts (115, 135)through the aligning arms (113, 133). Propellant from the propellanttank 230 is heated in the electrical heater 220 and exits the expansionnozzle 210 at a velocity in the range of about 7 to about 16 km/s topropel the launch vehicle 200 down the launch tube 100.

The propellant that is heated in the electrical heater 220 can comprisea light gas, and preferably is a gas that is ionizable at hightemperatures. In a preferred embodiment, the light gas used as thepropellant can be hydrogen. The electrical heater 220 preferably isadapted to heat the hydrogen or other propellant to a high temperature,such as in the range of about 1,000 K to about 100,000 K, about 2,000 Kto about 50,000 K, about 2,500 K to about 20,000 K, about 3,000 K toabout 15,000 K, about 3,500 K to about 10,000 K, or about 3,500 K toabout 5,000 K. In some embodiments, the exhausted gas exiting theexpansion nozzle 210 can be molecular hydrogen (i.e., with a molecularweight of 0.002 kg/mole). In further embodiments, the exhausted gasexiting the expansion nozzle 210 can be atomic hydrogen (i.e., with amolecular weight of 0.001 kg/mole). In still further embodiments, theexhausted gas exiting the expansion nozzle 210 can be hydrogen plasma(e.g., with a molecular weight of 0.0005 kg/mole).

In one exemplary embodiment, the electrical heater 220 can comprise aresistive heater such as illustrated in FIG. 11. The resistive heater1200 can comprise a resistive heater shell 1210 encasing an electricallyheated, heat cylinder 1220. The resistive heater shell may define acontainment vessel. The heat cylinder can be formed of a variety ofmaterials and composite structures. For example, a low density, highmelting point material such as carbon may be used. In some embodimentsthe cylinder may comprise carbon coated with a further material, such asdiamond, tungsten, hafnium carbide, or multiple layers of one or moredifferent materials. Such can be beneficial to improve heat transferperformance, strength, and reliability. The heat cylinder can be atranspiration tube element. For example, a porous tungsten heat cylindercan be used. The resistive heater shell 1210 can comprise any materialsuitable to contain the hot, expanding gas exiting the porous heatcylinder 1220 for controlled discharge through a gas discharge port1215. The propellant gas 5 enters the resistive heater 1200 through gasentry port 1213 through which it passes into the open core 1223. Insidethe porous heat cylinder 1220, the propellant gas 5 is heated to atemperature as described herein via electrical resistance heating fromthe electrical current passing through the electrical terminals (1203,1205). The heated gas expands (or transpires) outward through the poresin the porous heat cylinder walls 1221 and fills the expansion chamber1230 of the resistive heater 1200 prior to exiting the resistive heater1200 through the discharge port 1215.

In another exemplary embodiment, the electrical heater 220 can comprisean arc heater such as illustrated in FIG. 12. The arc heater 2200 cancomprise an arc heater shell 2210 encasing a swirl chamber 2230. The archeater shell may define a containment vessel and may comprisetranspiration cooled walls. The propellant gas 5 enters the arc heater2200 through gas entry port 2213 through which it passes into the swirlchamber 2230 wherein the propellant gas is heated to a temperature asdescribed herein via electrical arc passing between the electricalterminals (2203, 2205). As illustrated, the electrical terminals (2203,2205) of the arc heater 2200 can be coaxial and spaced apart by theswirl chamber 2230. In some embodiments, the electrical terminals can betranspiration cooled. The arc vortex within the swirl chamber 2230 canbe vortex stabilized. In particular, the propellant gas 5 is injectedtangentially into the swirl chamber via gas entry port 2213 rather thancoaxially with the gas discharge port 2215. This can form a helicalvortex as the fluid is heated by the arc discharge before being expandedthrough the gas discharge port 2215. Arc stability, heat transfer, andreliability may be improved by swirl stabilization, as well astranspiration, seeding, and like means.

Although hydrogen gas is a preferred propellant, other propellants maybe used, and various materials may be combined. For example, thepropellant gas may be seeded with an ionizable element and/or a furtherreactive element and/or an inert element. Non-limiting examples includecesium, rubidium, potassium, sodium, lithium, lithium hydride, argon,oxygen, and helium. The presence of such additional elements can beuseful to promote arc stability, conductivity, and ionization. In someembodiments, the seeded elements may be present as only a small fractionof the total mass of propellant, such as less than about 5%, less thanabout 4%, less than about 3%, less than about 2%, or less than about 1%by mass.

The expansion nozzle can take on any form suitable for expansion of thehot gas exiting the electrical heater so as to accelerate the launchvehicle in the manner described herein. In some embodiments, the exhaustnozzle can comprise a porous nozzle throat. Preferentially, the porescan be filled with a material that absorbs heat, such as by one or moreof melting, vaporization, and disassociation. In certain embodiments,the heat absorbing material can comprise solid hydrogen, solid lithium,or water ice. In further embodiments, the exhaust nozzle can include anozzle throat that is transpiration cooled, such as with a light gas,including hydrogen gas.

The propellant tank 230 utilized with the launch vehicle 200 may bereusable. Preferably, the propellant tank is sized to include asufficient volume of propellant (e.g., high pressure gas; liquid,semi-solid slush hydrogen, lithium hydride, water, or other materialsthat yield low molecular weight gases and high exhaust velocities uponbeing heated to high temperatures) to achieve exit of Earth's atmosphereor to substantially exit Earth's gravitational pull. In someembodiments, the propellant tank can be substantially cylindrical inshape. In some embodiments, the propellant tank can be formed of carboncomposite materials. The propellant tank particularly can be adapted tosupport the mechanical load of payload positioned in front of the tank,minus the pressurization between the tank and the payload, as furtherdiscussed below.

The propellant tank can be sized to have an outer diameter that issubstantially identical to the inner diameter of the launch tube. Insome embodiments, the propellant tank is in physical contact with theinner wall of the launch tube over a portion of the outer surface of thetank. In specific embodiments, the propellant tank may include slidingcontact strips on at least a portion of the outer surface. As such, themajority the tank structure is positioned slightly away from the tubeinner surface. The sliding contact strips can be adapted to vaporize asthe velocity of the launch vehicle increases and provide a low drag gasbearing to minimize frictional drag. The strips can be designed toproduce a vapor that is insulating so that it inhibits rather thanpromotes any arcing. For example, the sliding contact strips maycomprise pores filled with liquid sulfur hexafluoride.

The location of a launch system according to the present disclosure canvary. In some embodiments, the launch system can be located on theEarth. In other embodiments, the launch system can be at anon-terrestrial location, including in free space or on anothercelestial body.

During launch of the launch vehicle, the exit velocity can be in therange of about 2,000 to about 50,000 msec, about 4,000 to about 30,000msec, about 6,000 to about 15,000 msec, or about 8,000 to about 12,000msec. In some embodiments, the launch vehicle initially can beaccelerated to a velocity of about 100 to about 5,000 msec using analternate power means. For example, the initial launch velocity can beachieved using a single stage light gas gun. In such embodiments, thelight gas (e.g., hydrogen) can be preheated, particularly electricallypreheated, and more particularly preheated using electrical heating thatis derived from the same energy supply as the launch vehicle. Inalternate embodiments, the initial velocity achieved by such means canbe about 500 to about 3,000 msec or about 1500 to about 2500 msec.

As discussed above in relation to the propellant tank, the launchvehicle can be stabilized in one or more manners during passage thoughthe launch tube. One exemplary method is the use of the sliding contactstrips on the propellant tank. In other embodiments, the launchingmethod can be particularly important. For example, in certainembodiments, electrical heating is not utilized during the initiallaunch stage. As seen in FIG. 10, the launch tube 100 can comprise aninitial launch tube section 103. In this section, the launch vehicle canbe accelerated via a hot expanding gas (e.g., hydrogen). The utilizationof a light gas gun model at the initial launch stage can be useful toaccelerate the launch vehicle to as high a velocity as possible beforeelectrically power thrust is initiated. This can conserve propellantfrom the launch vehicle propellant tank and also conserve electricalenergy. This also can ensure that the sliding electrical contacts arealready moving at a high velocity before they begin conducting current.This can be particularly relevant when the sliding contacts are adaptedfor state transitions. The initial launch tube section (i.e., the firststage launcher) typically is not powered and is electrically isolatedfrom the second stage launch tube. This can be beneficial to avoid lowvelocity conduction where overheating may occur arising from excessivecontact duration of the sliding contacts at any given point on theconductive launch tube wall.

In some embodiments, the launch vehicle can be further stabilized withinthe launch tube via differential pressurization. As illustrated in FIG.13, shaded areas around the launch vehicle positioned within the launchtube can be differentially pressurized to minimize acceleration inducedmechanical stress on structures, particularly the nozzle, the heater,the propellant tank, and the payload.

Electrical heating of a low molecular weight gas, such as hydrogen, canprovide a uniquely high speed rocket exhaust as noted above that hasheretofore been unattainable with known chemical rocket technology. Thisin turn can lead to designs that can achieve, in exemplary embodiments,15% to 45% payload fractions to orbit. Accordingly, rather thanrequiring the use of a rocket having a mass on the order of 50 to 500tons, the launch systems of the present disclosure can launch packagesin a cost effective manner, the packages being orders of magnitudesmaller than rocket-based systems (e.g., 0.05 to 1 ton or 0.2 to 2tons).

The launch vehicle is accelerated inside an evacuated tube rather thanin free flight. The launch vehicle preferably can be disallowed fromexiting the launch tube unless the system confirms safe launchconditions exist. After exiting the launch tube, the launch vehicle canmaneuver through the atmosphere to orbit or to a specific destination(e.g., an extraorbital site in relation to space flight or a definedterrestrial location in relation to intra-atmospheric launches).

Returning to FIG. 1, the launch vehicle 200 also can comprise a payload240. The payload 240 can be removably connected to the propellant tank230. As illustrated, a payload connection element 250 provides theconnection, and any suitable means for removably connecting the payload240 to a further component of the launch vehicle 200 can be used. Thelaunch system 20 further can include a payload stabilizer 260, which cancomprise one or more arms or like element that is positioned between thepayload 240 and the inner tube wall 130 (preferably near the forward tipof the payload) and stabilizes the payload against radial movement whilepassing through the tube 100. The payload stabilizer 260 preferablydisengages from the payload at or near the exit 107 of the launch tube100.

The payload can be a container housing various types of cargo,including, but not limited to, human passengers, consumable resources,communication equipment, power components, arms, ordinances, rawmaterials, and the like. The nature of the cargo can, in someembodiments, define certain parameters of the launch system. Forexample, the dimensions of the launch tube and acceleration of thelaunch vehicle can be different for human passengers or cargo subject toadverse effects of experiencing excessive G forces. In some embodiments,the length of the launch tube 100 in meters as shown in FIG. 10(particularly the main launch tube section 105) can increase as the cuberoot of the launched mass in kilograms. In some embodiments, a launchtube has a length of up to about 500 to about 1000 miles in length.Further, launch conditions for humans, etc. can be limited such as toabout 2 G's to no more than 60 G's acceleration. In certain embodiments,the length of the launch tube 100 in meters can be equal to the squareof the launch velocity divided by twice the average acceleration of thelaunch.

Further considerations in relation to the launch package are describedbelow. In some embodiments, the launch package can have inertial sensorsand actuators that actively maintain its alignment and orientation whilebeing accelerated in the launch tube. In some embodiments, the launchpackage can be monitored during the launch acceleration interval forintegrity and nominal performance. Preferably, emergency procedures canbe implemented based on monitoring results to optimize the launch and toprotect the launch tube. Further, the launch can be aborted bydestroying the launch package immediately or shortly after its exit fromthe launch tube. In some embodiments, the launch package can beseparated from the remaining components of the launch vehicle during orimmediately after launch. These separate components can be defined as aflyout payload portion and a discarded or recycle portion. Separation ofcomponents can be significantly rapid and can utilize, for example, agas bag discharge or explosive bolt disconnects. The separation can beaided by the aerodynamic forces after exit. In particular embodiments,the flyout payload can have a heat shield with a transpiration coolednosetip to maintain the nosetip integrity, shape, sharpness, low drag,and low pressure moment during exit from the atmosphere. In someembodiments, the flyout payload can have a small positive stability,neutral stability, or a negative aerodynamic stability based on itscenter of pressure location relative to its center of mass location. Insome embodiments, the flyout payload can maneuver at high lateralacceleration levels to optimize flight path through the atmosphere andchange launch azimuth. In some embodiments, the flyout payload can havea high lift to drag ratio. In some embodiments, the payload can have alifting body design. In some embodiments, the flyout payload can haveaerodynamic control surfaces with very high speed response and low drag.In particular, the surfaces can be base split flaps or the surfaces canbe actuated with piezoelectric actuators.

In some embodiments, the flyout payload can be an orbital satellite. Forexample, the satellite can be a communications satellite, a sensorsatellite, resupply vehicle, or a weapon. In some embodiments, theflyout payload can be a suborbital payload. For example, the payload canbe a commercial package to be delivered rapidly to long distances, thepayload can be a sensor payload, the payload can be a UAV or otherunmanned vehicle, or the payload can be a weapon. In such embodiments,the payload may contain subparts that can be dispersed before impact,the payload can remain intact until impact, or multiple payloads canimpact at or near the same location for deep penetration. In someembodiments, the satellite can contain an inflatable solar array forpower. In some embodiments, the satellite can contain an inflatablemagnet array to effect attitude control in orbit. In some embodiments,the satellite can contain an inflatable antenna array to effectcommunications in orbit. In some embodiments, satellite containsinflatable structures to effect missions in orbit. In particular, theinflatable structures can harden to rigidity after deployment.

In some embodiments, the design lifetime of the satellite can be lessthan about 10 years, less than about 5 years, less than about 2 years,or less than about 1 year. In some embodiments, the satellite orbitalaltitude can be such that the orbital lifetime due to aerodynamic dragcan be less than about 5 years, less than about 2 years, less than about1 year, less than about 6 months, less than about 3 months, or less thanabout 1 month. In some embodiments, the satellite can achieve longerorbital lifetime through magnetic thrust against the Earth's magneticfield using an inflatable magnetic array, through pressure induced bysunlight and solar wind on an inflatable solar sail, or throughmagnetohydrodynamic (MHD) propulsion against ionized upper atmospheremolecules.

The launch system of the present disclosure can provide certainadvantages over known space launch systems. In some embodiments, payloadcost can be reduced through using commercial grade parts with highinitial failure rates and then iterating quickly through launch, fail,and redesign cycles to achieve higher and higher reliability quicklyover time. Further, the launcher and up to thousands of payloads can bedesigned simultaneously for a single purpose, if desired. In someembodiments, the payloads can be all communication satellites. In someembodiments, the satellites can be radiofrequency communicationsatellites. In some embodiments, the satellites can be opticalcommunications satellites. In some embodiments, the payloads can bereflective relays for millimeter waves or optical beams. In someembodiments, the payloads can be nuclear waste containers. In someembodiments, the flyout payload can have a heat shield with a porousnosetip filled with a material that absorbs heat by melting and/orvaporization and/or disassociation to maintain the nosetip integrity,shape, sharpness, low drag, and low pressure moment during exit from theatmosphere. In particular, the material can be solid hydrogen or lithiumor ice.

One embodiment of a payload for a launch vehicle according to thepresent disclosure is provided in FIG. 14 (in an atmospheric transitconfiguration) and FIG. 15 (in an on-orbit deployed configuration). Theembodiment illustrates an example integrating launch capabilities,satellite structure, and communications services. Specifically, FIG. 14,shows an external view of the exemplary launch vehicle payload componentas well as a cut-away view revealing an exemplary compartmentalizationof the multiple elements of the payload component. In the external view,the payload component has a conical shape to provide favorableaerodynamics, but other shapes are also encompassed. The payloadcomponent specifically is shown with aeromaneuver flaps and a nose coneheat shield. In the internal view, the payload component houses (frombase to tip) an orbital insertion motor, orbital insertion propellants,a solar cell array, attitude control, a communications payload, andavionics. The illustration of FIG. 15 shows the remaining aerobody ofthe payload component in connection with its inflated solar arrays,inflated attitude control system and inflated communications antenna.Further examples of a payload carrying a variety of useful elements fororbital delivery are provided in U.S. Pat. No. 6,921,051, the disclosureof which is incorporated herein by reference in its entirety.

In further embodiments, the present disclosure can provide methods forlaunching a payload. For example, in certain embodiments, the presentdisclosure can provide an electroantimagnetic launch method foraccelerating a launch vehicle. The method can comprise electricallyheating a propellant to form an expanding gas that accelerates thelaunch vehicle through a launch tube to a velocity of at least about5,000 with an acceleration force of about 2 to about 2,000 G's whilelimiting a magnetic field within the tube to no more than about 2Tesla.” The launch method can be defined by a variety of combinations ofthe further elements of the EAM launch system as otherwise describedherein.

EXAMPLE

Mathematical modeling of launch systems was carried out to exemplify theadvantages of the presently disclosed EAM launcher, particularly inrelation to typical, prior art EM launchers.

F=MA=0.5×L′×I ²

-   F=Force in Newtons-   M=Mass in kg-   A=acceleration in meters per second squared-   L′=the increase in inductance per meter of travel in the launch tube    in microhenries per meter-   I=current in amps

Modeling of a Typical Known Art Railgun (for a 1 Ton Payload)

-   L′=5 E-7 H/m-   I=20E6 A-   F=0.5×5E-7×(20E6)²=0.5×5E-7×400E12=1E8 N-   A=F/M=1E5 m/sec²=10,000 G's-   Typical efficiency demonstrated at 6000 m/sec is about 0.1% to 1%-   Typical efficiency demonstrated at 3000 m/sec is about 10-15%

Modeling of a Typical Known Art Coilgun (for a 1 Ton Payload)

-   L′=500 E-7 H/m-   I=2E6 A-   F=0.5×500E-7×(2E6)²=0.5×500E-7×4E12=1E8 N-   A=F/M=1E5 m/sec²=10,000 G's-   Highest velocity ever achieved by a coilgun was about 1000 m/sec-   Key problem is that the drive voltages required are:

Drive voltage=V=L′×I×Velocity

=500E-7×2E6×8000=400,000 volts.

It is believed that it has no previous work has heretofore achievesvoltages over about 50,000 volts in a coil launcher, and this haslimited the velocities that can be attained. Moreover, capacitors arethe only known power source to drive a coilgun. When considering anefficiency of 20%, one ton at 8800 m/sec requires 194 Gigajoules ofcapacitors. As capacitor power supplies presently cost roughly $1/Joule,this model would require $194B for the power supply alone.

EAM Launcher According to the Present Invention (for a 1 Ton Payload)

-   L′=0.2 E-7 H/m-   I=0.5 E6 A-   Magnetic force F=0.5×0.2E-7×(0.5E6)²=0.5×0.5E-7×0.25E12=6.25E3 N-   A=F/M=6.25 m/sec²=0.6 G's (so the magnetic “push” is 0.6 G's, versus    the “gas nozzle push” of around 500 G's).

As seen above, the present EAM launcher can lower required electricalcurrent by 40 fold versus the railgun and 4 fold versus the coilgun,thus reducing the resistive and arcing and magnetic energy storagelosses by 1600 times and 16 times respectively. The low launcher currentmakes it compatible with low cost power supplies.

The rocket propulsion effect has been demonstrated to achieve velocitiesover 20,000 msec in space. The presently disclosed EAM launcher isparticularly advantageous in light of the combination of an electricallypowered rocket in a conductive tube designed to maximize propulsionforce per unit current while eliminating magnetic fields and forces tothe maximum extent possible. This lowers losses due to resistive heatinglosses, arc losses, and stored magnetic energy losses.

Many modifications and other embodiments of the disclosure will come tomind to one skilled in the art to which this disclosure pertains havingthe benefit of the teachings presented in the foregoing descriptions andthe associated drawings. Therefore, it is to be understood that thedisclosure is not to be limited to the specific embodiments disclosedherein and that modifications and other embodiments are intended to beincluded within the scope of the appended claims. Although specificterms are employed herein, they are used in a generic and descriptivesense only and not for purposes of limitation.

1. A launch vehicle adapted for high velocity delivery of a payload, thelaunch vehicle comprising: a payload container; a propellant tankcontaining a propellant; an electrical heater in fluid connection withthe propellant tank and adapted for electrical heating of the propellantto form an exiting exhaust; and one or more electrical contacts adaptedfor directing flow of electrical current through the electrical heater.2. The launch vehicle according to claim 1, further comprising anexpansion nozzle in fluid communication with the exiting exhaust fromthe electrical heater.
 3. The launch vehicle according to claim 1,wherein the electrical heater is a resistive heater.
 4. The launchvehicle according to claim 3, wherein the resistive heater comprises anelectrically heated porous cylinder inside a containment vessel.
 5. Thelaunch vehicle according to claim 4, wherein the electrically heatedporous cylinder comprises tungsten walls.
 6. The launch vehicleaccording to claim 1, wherein the electrical heater is an arc heater. 7.The launch vehicle according to claim 6, wherein the arc is a swirlstabilized vortex arc.
 8. The launch vehicle according to claim 7,wherein the arc heater comprises a swirl chamber inside a containmentvessel.
 9. The launch vehicle according to claim 8, wherein the archeater comprises coaxial electrical terminals spaced apart by the swirlchamber.
 10. The launch vehicle according to claim 1, wherein theelectrical contacts comprise sliding electrical contacts.
 11. A launchsystem comprising: a launch vehicle according to claim 1; a launch tubecomprising two or more concentric, electrically conductive tubesseparated by an insulator, the launch tube being adapted for propulsionof the launch vehicle therethrough; and an electrical energy source. 12.The launch system according to claim 11, wherein the electrical energysource comprises a battery bank.
 13. The launch system according toclaim 12, wherein the electrical energy source further comprises aninductor.
 14. A method for launching a payload, the method comprising:providing the payload in a payload container of a launch vehicleincluded in the launch system according to claim 11; and electricallyheating the propellant in the electrical heater to form the exitingexhaust at a velocity sufficient to propel the payload out of the launchtube.
 15. An electro antimagnetic launch method for accelerating alaunch vehicle to a high velocity, the method comprising electricallyheating a propellant to form an expanding gas that accelerates thelaunch vehicle through a launch tube to a velocity of at least about5,000 with an acceleration force of about 2 to about 2,000 G's whilelimiting a magnetic field within the tube to no more than about 2 Tesla.